XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY LS(1)-0421MOD AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -16.500 -0.7468 0.09044 0.08318 -0.0610 1.0000 0.0486 -16.250 -0.7727 0.08281 0.07528 -0.0651 1.0000 0.0488 -16.000 -0.7893 0.07705 0.06930 -0.0678 1.0000 0.0492 -15.750 -0.8015 0.07222 0.06426 -0.0698 1.0000 0.0497 -15.500 -0.8110 0.06795 0.05978 -0.0713 1.0000 0.0503 -15.250 -0.8190 0.06407 0.05566 -0.0723 1.0000 0.0510 -15.000 -0.8213 0.06100 0.05243 -0.0728 1.0000 0.0517 -14.750 -0.8137 0.05909 0.05054 -0.0729 1.0000 0.0524 -14.500 -0.8083 0.05703 0.04843 -0.0730 1.0000 0.0532 -14.250 -0.8034 0.05492 0.04626 -0.0730 1.0000 0.0541 -14.000 -0.7984 0.05283 0.04406 -0.0728 1.0000 0.0552 -13.750 -0.7925 0.05082 0.04190 -0.0725 1.0000 0.0563 -13.500 -0.7842 0.04908 0.04008 -0.0720 1.0000 0.0573 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0.9830 0.02619 0.01710 -0.0543 0.4033 0.7911 6.000 1.0076 0.02648 0.01735 -0.0545 0.3923 0.7941 6.250 1.0244 0.02682 0.01774 -0.0529 0.3817 0.7963 6.500 1.0416 0.02716 0.01805 -0.0514 0.3716 0.7989 6.750 1.0570 0.02762 0.01857 -0.0498 0.3604 0.8021 7.000 1.0732 0.02807 0.01896 -0.0485 0.3503 0.8057 7.250 1.0872 0.02863 0.01959 -0.0471 0.3388 0.8095 7.500 1.1039 0.02929 0.02023 -0.0465 0.3278 0.8130 7.750 1.1147 0.02999 0.02095 -0.0445 0.3169 0.8154 8.000 1.1261 0.03080 0.02180 -0.0429 0.3062 0.8181 8.500 1.1494 0.03275 0.02376 -0.0405 0.2847 0.8244 8.750 1.1615 0.03389 0.02488 -0.0398 0.2747 0.8280 9.000 1.1745 0.03518 0.02616 -0.0395 0.2644 0.8318 9.250 1.1818 0.03641 0.02742 -0.0379 0.2557 0.8347 9.500 1.1890 0.03778 0.02880 -0.0367 0.2470 0.8379 9.750 1.1978 0.03923 0.03028 -0.0358 0.2389 0.8413 10.000 1.2069 0.04081 0.03186 -0.0353 0.2310 0.8447 10.250 1.2172 0.04245 0.03350 -0.0352 0.2239 0.8480 10.500 1.2255 0.04416 0.03527 -0.0347 0.2168 0.8509 10.750 1.2312 0.04581 0.03687 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0.1404 0.9285 15.750 1.3328 0.09131 0.08372 -0.0339 0.1380 0.9339 16.000 1.3304 0.09489 0.08744 -0.0349 0.1356 0.9396 16.250 1.3309 0.09775 0.09039 -0.0354 0.1335 0.9467 16.500 1.3386 0.09962 0.09227 -0.0357 0.1317 0.9560 16.750 1.3460 0.10124 0.09388 -0.0359 0.1300 0.9745 17.000 1.3224 0.10829 0.10128 -0.0392 0.1279 1.0000 17.250 1.3014 0.11571 0.10895 -0.0430 0.1256 1.0000 17.500 1.2886 0.12195 0.11537 -0.0465 0.1234 1.0000 17.750 1.2937 0.12522 0.11867 -0.0485 0.1216 1.0000 18.000 1.3153 0.12563 0.11902 -0.0491 0.1201 1.0000